Gas turbine engine with heat exchanger diagnostics

ABSTRACT

A gas turbine engine has a compressor section and a turbine section. A secondary cooling air includes a first fluid connection to tap cooling air and pass the cooling air through a plurality of tubes, and a second fluid connection for returning air from the tubes back to at least one of the compressor and turbine for cooling. A sensor senses a condition of the cooling air downstream of the tubes and a control compares the sensed condition of the cooling air to an expected condition, and to identify a potential concern in the cooling air system should the sensed condition differ from the expected condition by more than a predetermined amount.

BACKGROUND OF THE INVENTION

This application relates to a system for monitoring the operation of aheat exchanger providing cooling air in a gas turbine engine.

Gas turbine engines are known and typically include a fan delivering airinto a bypass duct as propulsion and delivering air into a compressor ascore engine air. The air is compressed and then delivered into acombustor where it is mixed with fuel and ignited. Products of thiscombustion pass downstream over turbine rotors driving them to rotate.

As known, the products of combustion are hot. Further, the compressorcomponents can become hot, particularly, near downstream locations.

As such, cooling air is provided to components within the gas turbineengine.

Historically, the fan rotated at the same speed as a turbine rotor. Thislimited the design of the gas turbine engine as it would be desirablefor the turbine rotor to rotate at faster speeds and the fan rotor torotate at slower speeds. More recently, a gear reduction has beenlocated between the turbine rotor and the fan rotor.

With the inclusion of the gear reduction, the temperatures experiencedby the compressor and turbine sections has increased dramatically.

SUMMARY OF THE INVENTION

In a featured embodiment, a gas turbine engine has a compressor sectionand a turbine section. A secondary cooling air includes a first fluidconnection to tap cooling air and pass the cooling air through aplurality of tubes, and a second fluid connection for returning air fromthe tubes back to at least one of the compressor and turbine forcooling. A sensor senses a condition of the cooling air downstream ofthe tubes and a control compares the sensed condition of the cooling airto an expected condition, and to identify a potential concern in thecooling air system should the sensed condition differ from the expectedcondition by more than a predetermined amount.

In another embodiment according to the previous embodiment, thecondition is a pressure of the cooling air downstream of the tubes froma heat exchanger.

In another embodiment according to any of the previous embodiments, thefirst and second connections are fluid conduits.

In another embodiment according to any of the previous embodiments, thepressure is sensed downstream of the second fluid connection.

In another embodiment according to any of the previous embodiments, thecooling air is returned into a diffuser case downstream of thecompressor section for transfer to at least one of the turbine and thecompressor sections.

In another embodiment according to any of the previous embodiments, thesensor senses the pressure within a part of the diffuser case which isdedicated to passing cooling flow to at least one of the turbine andcompressor sections.

In another embodiment according to any of the previous embodiments, thesensor is positioned on the diffuser.

In another embodiment according to any of the previous embodiments, apressure tap taps pressure from within the diffuser to the sensor, whichis remote from the diffuser.

In another embodiment according to any of the previous embodiments, thesensor is located to be line replaceable.

In another embodiment according to any of the previous embodiments, areference pressure is also sensed and the sensed cooling air pressure iscompared to the sensed reference pressure at the expected condition.

In another embodiment according to any of the previous embodiments, thereference pressure is a pressure sensed downstream of the compressorsection.

In another embodiment according to any of the previous embodiments, theexpected condition is an artificially determined reference pressure.

In another embodiment according to any of the previous embodiments, thecomparison may result in varying levels of indicated maintenance.

In another embodiment according to any of the previous embodiments, ifthe comparison results in the sensed pressure differing from theexpected pressure by a first amount, then routine maintenance may beindicated whereas if the sensed pressure differs from the expectedpressure by a second greater amount, a step more than routinemaintenance may be indicated.

In another embodiment according to any of the previous embodiments, themore drastic step may be a power reduction.

In another embodiment according to any of the previous embodiments, thecondition is sensed downstream of the second fluid connection.

In another embodiment according to any of the previous embodiments, areference condition is also sensed and the sensed cooling air conditionis compared to the sensed reference condition as the expected condition.

In another embodiment according to any of the previous embodiments, theexpected condition is an artificially determined reference condition.

In another featured embodiment, a method includes tapping compressedcooling air and passing the cooling air through a heat exchanger, andreturning air from the heat exchanger back to at least one of acompressor and turbine in a gas turbine engine for cooling. A conditionof the cooling air downstream of the heat exchanger is sensed andcompared to an expected condition. A potential concern in the coolingair system is identified should the sensed condition differ from theexpected condition by more than a predetermined amount.

In another embodiment according to the previous embodiment, thecondition is a pressure of the cooling air downstream of the heatexchanger.

These and other features may be best understood from the followingdrawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2A shows a first gas turbine engine arrangement.

FIG. 2B shows alternative locations.

FIG. 3 shows a detail.

FIG. 4 is a flowchart.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

FIG. 2A shows an engine 80 incorporating a fan rotor 82 within a fancase 84. An intermediate engine case 86 surrounds a low pressurecompressor 88 rotating with a low pressure turbine 90. A gear reduction91 drives the fan rotor 82 through the turbine 90.

A high pressure compressor 92 rotates with a high pressure turbine 94. Acombustor 96 is shown schematically. A diffuser 98 is positioneddownstream of the high pressure compressor 92. A heat exchanger 100 isshown tapping high pressure air 102. That air is cooled by bypass air inbypass duct 83 and is returned through conduit 104 to the diffuser 98.When the air is returned through conduit 104, it passes through diffuser98, optionally cooling downstream locations in the high pressurecompressor 92 such as by diverting a portion of bypass air or other airthat is cooler than the compressor discharge air in an upstreamdirection to a last stage of the high pressure compressor 92, and thenpasses as shown at 107 to the high pressure turbine 94 for cooling.

As mentioned above, the cooling loads on this air increase dramaticallywith recent developments in gas turbine engines.

FIG. 2B shows alternative engine 180. In engine 180, the intermediatecase 186 is again positioned outwardly of a low pressure compressor 188.Tap 190 taps air for cooling from low pressure compressor 188, andthrough a boost compressor 192 to a heat exchanger 194. A scoop inlet196 takes bypass air from bypass duct 183 and passes it over the heatexchanger 194 to an outlet 198. Air from tap 190, having been cooled inheat exchanger 194, passes through conduit 200 back into the engine forcooling such as shown in the FIG. 2A embodiment.

It should be understood that either location for the heat exchangercould be utilized in combination with either location for the tap forthe cooling air. In addition, the cooling air could be tapped fromupstream locations in the high pressure compressor 92.

The teachings of this disclosure could apply to any combination of taplocation, return location, and heat exchanger location. With such anarrangement, monitoring the viability of the heat exchanger becomesimportant. Applicant has recognized that it would be undesirable forthere to be a failure in the supply of cooling air, particularly, withmodern high temperature engines.

Thus, FIG. 3 shows an arrangement wherein the diffuser 98, which ispositioned downstream of the last stage 106 of the high pressurecompressor 92, receives the cooling air through an inlet 108. Theconduit 104 is shown connected to the opening 108. A tap 110 is shown totap the pressure within the diffuser 98. The tap 110 will “see” thepressure delivered by the conduit 104 into the opening 108. The tap 110includes a sensor 114 at some location, which is connected as shown at112 to a control 119. The sensor 114 be located to be within thediffuser 98 or remote from the diffuser 98, but connected by a tube. Thecontrol 119 may be a stand-alone control or may be part of a fullauthority digital engine control (FADEC).

Another sensor 115 is illustrated and may sense a pressure, such as thedischarge pressure downstream of the compressor 106. Other locations mayalso be sensed.

A designer may know that the pressure delivered to conduit 104,downstream of the heat exchanger 100 or 194, should approximate thepressure sensed by the sensor 115. If the two pressures are compared anddiffer by more than a predetermined amount, then a flag may be set asshown schematically at 117, which is indicative of a potential need formaintenance. Varying degrees of flags may be set. As an example, if thepressures are within 95 percent of each other, then maintenance may beset within a period of time. On the other hand, if the pressure sensedby sensor 114 is less than 90 percent of the pressure sensed by sensor150, then indication may be set to flag that a power reduction isindicated. That is, with a greater difference, more drastic correctivesteps may be indicated.

By sensing the pressure downstream of the heat exchanger 100 or 194, thesystem will be able to identify a failure in the heat exchanger 100 or194, or any of the conduits 102/104/190/200. As such, the system ensuresproper operation of the cooling air system and the supply of cooling airto the components requiring cooling. The heat exchanger could be anynumber of configurations, including a plurality of tubes for containingthe cooling air.

As known, the diffuser typically includes a plurality ofcircumferentially spaced vanes and the air passes through thesecircumferentially spaced vanes. There may be a plurality of sensors in aplurality of these vanes to provide redundant information.

While a particular comparison is disclosed, it should be understood thatthe pressure sensed by sensor 114 could instead simply be compared tosome preset or predetermined limit. This preset or predetermined limitcould be based upon engine operating conditions. In addition, while aparticular location is shown for sensor 115, other locations could beutilized which may be at other pressures. As an example, the referencepressure could come from a more intermediate location in the compressorwith the pressure sensed by sensor 114 being expected to be higher thanthe referenced pressure by a predetermined amount.

Broadly speaking, a system could be said to include a gas turbine enginewith a compressor section and a turbine section. A cooling air systemincludes a fluid connection to tap air 102/190 and pass the cooling airthrough a heat exchanger 100/194. A fluid connection returns air fromthe heat exchanger back to at least one of the compressor and turbinefor cooling. A sensor 114 senses a condition of the cooling airdownstream of the heat exchanger and compares the condition of thecooling air to an expected condition. A control identifies a potentialconcern in the cooling air system should the sensed condition differfrom the expected by more than a predetermined amount.

A flowchart is shown in FIG. 4. At step 116, a cooling air pressure issensed. At step 118, it is compared to an expected pressure. Asexplained above, this expected pressure may be a preset limit, or alimit determined artificially or may be a sensed reference pressureelsewhere in the engine.

If the sensed cooling air pressure compares to the expected closely,then the method returns to step 116. On the other hand, if the twopressures are different, or not as expected, then a flag 122 is set.

While a pressure is sensed, other conditions, such as temperature of thecooling air may be sensed as indicative of the condition of the coolingair system and, in particular, the heat exchanger and its plumbingconnections.

In the location illustrated in FIG. 3, the sensor 114 is “linereplaceable.” That means it can be replaced without disassembly of thecompressor module. This is a beneficial location, as requiringdisassembly of a compressor module would make replacement of the sensoran expensive step.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

The invention claimed is:
 1. A gas turbine engine comprising: acompressor section and a turbine section, said compressor sectionincluding a low pressure compressor and a high pressure compressor, andsaid turbine section including a low pressure turbine and a highpressure turbine; a fan rotor, said low pressure turbine driving saidlow pressure compressor, and said low pressure turbine driving said fanrotor through a gear reduction such that said fan rotor rotates at alower speed than said low pressure turbine; said high pressure turbinedriving said high pressure compressor; a cooling air system including afirst fluid connection to tap cooling air and pass the cooling airthrough a heat exchanger, and a second fluid connection for returningair from the heat exchanger back to at least one of the compressorsection and the turbine section for cooling; and a first sensor forsensing a pressure of the cooling air downstream of the heat exchangerand a control to compare the sensed pressure of the cooling air to anexpected pressure, and said control is programmed to identify acondition of the cooling air system when the sensed pressure differsfrom the expected pressure by more than a predetermined amount, saidcondition being indicative of a failure in said heat exchanger, saidfirst fluid connection, or said second fluid connection, and saidcontrol setting a maintenance flag such that when cooling air is notproperly passing to the at least one of the compressor section and theturbine section the maintenance flag is set; and wherein said firstfluid connection and said second fluid connection are fluid conduits. 2.The gas turbine engine as set forth in claim 1, wherein said pressure issensed downstream of said second fluid connection.
 3. The gas turbineengine as set forth in claim 2, wherein said cooling air is returnedinto a diffuser downstream of said compressor section for transfer to atleast one of the turbine section and the compressor section.
 4. The gasturbine engine as set forth in claim 3, wherein said first sensor sensesthe pressure within a part of said diffuser which is connected to passcooling air to at least one of the turbine section and the compressorsection.
 5. The gas turbine engine as set forth in claim 4, wherein saidfirst sensor is positioned on said diffuser.
 6. The gas turbine engineas set forth in claim 4, wherein a pressure tap taps air from withinsaid diffuser to said first sensor, said first sensor being remote fromsaid diffuser.
 7. The gas turbine engine as set forth in claim 1,wherein said first sensor is located to be line replaceable.
 8. The gasturbine engine as set forth in claim 1, wherein a reference pressure isalso sensed and said pressure sensed by said first sensor is compared tosaid reference pressure as said expected condition.
 9. The gas turbineengine as set forth in claim 8, wherein said reference pressure is apressure sensed downstream of said compressor section.
 10. The gasturbine engine as set forth in claim 1, wherein said expected conditionis an artificially determined reference pressure.
 11. The gas turbineengine as set forth in claim 1, wherein said control is programmed suchthat the comparison results in varying levels of the maintenance flag.12. The gas turbine engine as set forth in claim 11, wherein saidcontrol is programmed such that when the comparison results in thepressure sensed by the first sensor differing from said expectedpressure by a first amount, then the maintenance flag indicates routinemaintenance whereas when said pressure sensed by said first sensordiffers from said expected pressure by a second greater amount, then themaintenance flag indicates a more drastic step than routine maintenance.13. The gas turbine engine as set forth in claim 12, wherein said moredrastic step is a power reduction of the gas turbine engine.
 14. A gasturbine engine comprising: a compressor section and a turbine section; acooling air system including a first fluid connection to tap cooling airand pass the cooling air through a heat exchanger, and a second fluidconnection for returning air from the heat exchanger back to at leastone of the compressor section and the turbine section for cooling; afirst sensor for sensing a pressure of the cooling air downstream of theheat exchanger and a control to compare the sensed pressure of thecooling air to an expected pressure, and said control is programmed toidentify a condition of the cooling air system when the sensed pressurediffers from the expected pressure by more than a predetermined amount;and said control is programmed such that condition being indicative of afailure in said heat exchanger, said first fluid connection, or saidsecond fluid connection, and said control setting a maintenance flagsuch that when cooling air is not properly passing to the at least oneof the compressor section and the turbine section the maintenance flagis set; wherein a reference pressure is also sensed and said pressuresensed by said first sensor is compared by said control to saidreference pressure as said expected pressure; wherein said referencepressure is a pressure sensed downstream of said compressor section;wherein said control is programmed such that said comparison results invarying levels of the maintenance flag; and wherein said controlprogrammed such that when the comparison results in the pressure sensedby the first sensor differing from said expected pressure by a firstamount, then the maintenance flag indicates routine maintenance whereaswhen said pressure sensed by said first sensor differs from saidexpected pressure by a second greater amount, then the maintenance flagindicates a more drastic step than routine maintenance.
 15. The gasturbine engine as set forth in claim 14, wherein said expected pressureis an artificially determined reference pressure.
 16. The gas turbineengine as set forth in claim 14, wherein said pressure is senseddownstream of said second fluid connection.
 17. The gas turbine engineas set forth in claim 14, wherein said more drastic step is a powerreduction of the gas turbine engine.
 18. A gas turbine enginecomprising: a compressor section and a turbine section; a cooling airsystem including a first fluid connection to tap cooling air and passthe cooling air through a heat exchanger, and a second fluid connectionfor returning air from the heat exchanger back to at least one of thecompressor section and the turbine section for cooling; and a firstsensor for sensing a pressure of the cooling air downstream of the heatexchanger and a control to compare the sensed pressure of the coolingair to an expected pressure, and said control is programmed to identifya condition of the cooling air system when the sensed pressure differsfrom the expected pressure by more than a predetermined amount, saidcondition being indicative of a failure in said heat exchanger, saidfirst fluid connection, or said second fluid connection, and saidcontrol setting a maintenance flag such that when cooling air is notproperly passing to the at least one of the compressor section and theturbine section the maintenance flag is set, wherein when a differencebetween said sensed pressure and said expected pressure exceeds amaximum, said maintenance flag includes a reduction to of power of thegas turbine engine.